The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor and ignited for generating hot combustion gases. Energy is extracted from the combustion gases in turbines disposed downstream therefrom. A high pressure turbine powers the compressor, and a low pressure turbine powers a fan in a typical turbofan aircraft engine application.
Each turbine stage includes a stationary turbine nozzle having a row of vanes which direct the combustion gases to a cooperating row of turbine rotor blades. The vanes and blades are typically hollow and provided with air bled from the compressor for cooling the vanes and blades during operation.
Turbine vane and blade cooling art is quite crowded with a myriad of cooling configurations found therein specifically configured for cooling the various portions of the airfoils defining the vanes and blades. Each airfoil has a generally concave pressure side and an opposite generally convex suction side extending axially between leading and trailing edges and radially in span from an inner root to an outer tip.
In view of the three dimensional complex combustion gas flow distribution over the airfoils, the different portions thereof are subjected to different heat loads during operation. The heat, in turn, generates thermal stress in the airfoils which must be suitably limited for prolonging the life of the airfoil.
The airfoils are typically manufactured from superalloy cobalt or nickel based materials having sustained, strength under high temperature operation. The useful life of the airfoils is limited by the maximum stress experienced therein irrespective of its particular location in the airfoil.
Accordingly, the prior art includes various forms of internal cooling channels having various forms of heat transfer increasing turbulator ribs or pins therein for cooling the various portions of the airfoil with different effectiveness.
For example, U.S. Pat. No. 6,174,134xe2x80x94Lee et al, assigned to the present assignee, discloses a multiple impingement airfoil cooling configuration for effecting enhanced cooling in a trailing edge region of a turbine blade. However, that turbine airfoil is specifically configured for a relatively large turbofan engine and, correspondingly, the turbine airfoil itself is relatively large. The trailing edge cooling configuration disclosed in this patent has particular utility in the large airfoil size, with Reynolds numbers greater than about 30,000 for the impingement air directed against the pressure side trailing edge turbulators disclosed therein.
The assignee is developing another, smaller gas turbine engine having correspondingly smaller turbine blades which are not amenable to the cooling configuration disclosed in the Lee et al patent. Turbine airfoil cooling features are not readily scaled down in size from large turbine blades to small turbine blades in view of the inherent nature of heat transfer characteristics.
For example, attempting to scale down the configuration of the Lee et al patent in a smaller turbine airfoil would result in a Reynolds number for the impingement cooling air in a trailing edge cavity of substantially less than the 30,000 effected in the large airfoil. Correspondingly, insufficient heat transfer would be available for adequately cooling the small airfoil using the configuration of the larger Lee et al large airfoil.
Accordingly, it is desired to provide an improved impingement cooling configuration for the trailing edge region of relatively small turbine blade airfoils.
A turbine airfoil includes pressure and suction sidewalls having first and second flow channels disposed therebetween and separated by a longitudinally extending bridge. The bridge includes a row of inlet holes, and a row of outlet holes extends from the second channel toward the trailing edge of the airfoil. A row of turbulator ribs is disposed inside the second channel along the pressure sidewall and are longitudinally elongate and substantially colinear. The ribs face the inlet holes for crossover impingement cooling from the air channeled therefrom.